This invention relates to rotor blade systems and rotor aircraft, such as helicopters, in general, and in particular, to a mechanism that enables the pitch of each of the blades of the main rotor of the aircraft to be controlled individually and independently of the others.
As illustrated in the partial perspective view of FIG. 1, most rotor aircraft, such as helicopters, include one or more power driven main rotors 1 that are equipped with a hub 2 rotatably supported on a rotor mast 5 and having a plurality of elongated blades 3 extending radially outward from it, each of which has an airfoil cross-section (not illustrated) and is coupled to the hub at the inner end thereof by three hinges 4 that enable the blade to rotate about three axes relative to the hub, i.e., to “flap” up and down, to “lag” forward and aft, and to “pitch” up and down and thereby change its angle of attack relative to the stream of air moving past it.
To move the aircraft vertically, the respective pitches of the rotating blades are all changed simultaneously, or “collectively,” which is effected by a “collective pitch lever” coupled to the blades through a rotating “swash plate” that is coupled to the blades by respective linkages.
On the other hand, horizontal movement of the aircraft is achieved by tilting the rotor such that the thrust of the rotor resolves into two components, a “lift” component that supports the weight of the aircraft, and a “horizontal thrust” component that propels the aircraft horizontally in the desired direction. This tilting of the rotor is effected by tilting the swash plate, which results in a “cyclic pitch control” of the blades, in which the pitch of each of the blades changes twice, i.e., one pitch cycle, per revolution of the hub. For example, to move the aircraft directly forward, the pitch, or angle of attack, of each blade is increased each time that blade passes over the tail of the aircraft, such that the lift developed by that blade is then temporarily greater than that of the other blades, and thereby results in a forward thrust component being applied to the aircraft by the rotor.
As a result of the foregoing method of operation and the effect of the relative speed of the aircraft moving through air, conventional rotor aircraft have a limited forward air speed, viz., about 180 knots (˜207 mph), due to the blade tip speed approaching the speed of sound on the advancing blade, and a stall condition occurring on the retreating blade. Additionally, when these limiting conditions are being approached, large vibrations begin to occur in the rotor, which causes component fatigue and increased pilot mental and physical work load. The vibrations cause the entire vehicle, including the pilot and aircraft cockpit controls, to shake and the aircraft displays to become blurred.
One effort to address the foregoing problem has been the development of so-called “tilt rotor” aircraft to provide enhanced helicopter lift capability, higher forward airspeed and reduced vibration. However, this approach adds wings, aero-surface controls, complex rotor conversion mechanisms, weight and cost.
In another approach, rotor aircraft designs have been developed implementing so-called “compound” systems that have both conventional rotor systems and additional forward propulsion systems. These compound designs typically also include additional lifting surfaces and aero-surface controls, which add significant complexity, weight and cost to the aircraft.
In a third approach, individual blade control (IBC) is used in conjunction with a lower rotor speed and “reverse rotor flow” technology. IBC systems enable the direction of pitch of each blade to be varied independently of the others and more than twice per revolution of the hub, as occurs in conventional rotor aircraft. EBC also enables the rotor system track and balance procedure to be implemented in software, thereby eliminating the time consuming process of manually adjusting the length of each pitch link. Typical approaches to IBC utilize either electrical motor actuators and slip rings, or hydraulic actuators, hydraulic swivels and electrical slip rings. Both approaches are complicated, add extensive installation congestion in the rotating section of the rotor system, and significantly reduce control reliability of the “flight critical” rotor system. The electric approach utilizes a screw mechanism that is susceptible to jamming and is dependent on slip ring technology, which is unreliable. Furthermore, lightening strike attachment to the rotor hub is a common occurrence and may completely eliminate all electrical control. The hydraulic approach is dependant on both electrical and hydraulic slip ring technology, neither of which is reliable. The lightening strike problem also exists with this design. External hydraulic leaks are centrifugally distributed onto numerous aircraft components, including the exterior body and windshield and require extensive clean up. The mass of the rotating power control actuators also introduces new stresses into the flight critical rotor system.
In light of the foregoing problems, there is a long felt but as yet unsatisfied need in the field of rotor aircraft for a simpler, substantially more reliable, less expensive and lighter weight mechanism for providing individual blade control (IBC) for the rotor(s) of a rotor aircraft.